This invention relates generally to methods and apparatus for increasing efficiency of airfoils, and more particularly to methods and apparatus for extracting from boundary layers of airfoils.
A gas turbine engine such as that configured for powering an aircraft in flight conventionally includes in serial flow communication a fan, a compressor, a combustor, a high pressure turbine (HPT), and a low pressure or power turbine (LPT). Ambient air enters the fan wherein it is initially pressurized, and in turn a portion thereof flows to the compressor wherein it is further pressurized and discharged to the combustor wherein it is mixed with fuel and ignited for generating hot combustion gases which flow downstream to the HPT. The HPT includes one or more stages of turbine blades specifically configured for extracting energy from the combustion gases for powering the compressor through a shaft connected therebetween. The combustion gases lose pressure in the HPT and then flow to the LPT which includes additional turbine blades also configured for extracting additional energy from the lower pressure combustion gases for powering the fan connected thereto by another shaft.
The fan and compressor include respective rotor blades which are configured for pressurizing the relatively cool air which is in contrast to the turbine blades of the HPT and the LPT which are configured for extracting energy from the hot combustion gases with a resulting reduction in pressure thereof. The energy extracted from the combustion gases is in turn imparted to the air being pressurized in the fan and compressor.
Rotor blades, fan blades, and compressor blades all represent types of airfoils. Both fan blades and compressor blades are effective for imparting energy into the air for increasing its pressure to different levels. Fan blades are relatively large for moving larger amounts of airflow at reduced pressure for providing a substantial portion of propulsion thrust from an engine. Fan blades are typically configured in one or two stages for use in conventional high bypass, turbofan, commercial aircraft engines or lower bypass military engines.
Rotor blades found in a typical axial compressor are configured in a substantial number of axial stages with each succeeding stage having smaller and smaller rotor blades for incrementally increasing pressure of the airflow channeled therethrough.
A large portion of the aerodynamic losses of transonic rotor blades and most of the aerodynamic losses of subsonic rotor blades are localized in boundary layers around the blade and the hub flowpath.
In at least one known configuration, blade surface, hub contour boundary layers, and tip clearance leakage flows develop without interruption to derate and limit potential performance and aerodynamic stability of compressive rotor blading. Incurred losses are passed on to downstream blading in the form of wakes and vortices that interact with the downstream blading to create further losses, possible aerodynamic instabilities, and noise.
U.S. Pat. No. 5,480,284 to Wadia, et al. describes a self-bleeding rotor blade and method of operation for reducing boundary layer thickness for improved performance. The rotor blade includes a suction surface configured for pressurizing air flowable thereover with bleed apertures being disposed therein for bleeding a portion of the boundary layer air from the suction surface during operation and thereby decreasing its thickness for improving aerodynamic performance of the blade.